Satellite module for attitude determination

This application is a U.S. National Phase Application of International Application No. PCT/IB2020/056693 filed on Jul. 16, 2020, which claims priority to Italian Application No. 102019000012498 filed on Jul. 22, 2019, the disclosures of which are incorporated by reference herein in their entirety.

The present invention relates to a satellite module for attitude determination, i.e. for determining how a satellite is oriented in space. The attitude determination module (AD) can be used as part of an Attitude Determination and Control System (ADCS) in a satellite.

ADCS systems determine the attitude of a satellite in order to set and/or correct the attitude in such a way that the satellite is able to accomplish the mission for which it was placed in orbit or launched into space.

The AD attitude determination is usually entrusted to sensors and processors capable of measuring and processing specific physical quantities to process accurate information on the spatial orientation of the satellite with respect to a relative reference system (for example integral with the satellite or to the earth) or to a system of inertial reference (for example integral with the stars).

The sensors typically used for the attitude determination are magnetometers that measure the intensity and direction of the earth's magnetic field, gyroscopes, sun sensors that identify the direction and/or position of the sun, stellar sensors that identify and recognize a set of stars, earth sensors that identify a direction pointing towards the centre of the earth and horizon sensors that identify the earth's horizon.

By combining together the information coming from two or more of said sensors (preferably from at least three of said sensors) it is possible to determine the orientation in space of the satellite.

The sensors are mounted on board the satellite so that they can perform their function during the operating life or mission of the satellite.

Since each type of sensor is designed to detect a specific quantity, each sensor is usually a stand alone sensor, that is to say it is a sensor capable of operating independently of the other sensors, and requires a dedicated connection for interacting with the ADCS system both in terms of electrical power supply (when necessary for sensor operation) and in terms of data transmission.

Each type of sensor is installed on the satellite in the most suitable position for detecting, during the mission, the specific information for which it was designed.

The Applicant has noted that the positioning of the sensors of the AD system is therefore very critical and very often needs to previously know the details of the mission to be carried out by the satellite as well as the attitude that the satellite will have to assume during the mission.

By way of example, the Applicant has verified that since the terrestrial globe has an angular diameter of about 85 degrees at a distance of about 500 Km from the earth's surface, the stellar sensor (typically a camera or a tele-camera) typically has a very small field of view (FOV), for example having an angular diameter comprised between 4 degrees and 7 degrees, in order to guarantee to frame a portion of stars, however small, not influenced by the terrestrial globe, the sun or the terrestrial albedo.

Given the limited field of view of the stellar sensor, it is necessary to carefully choose the portion of the satellite on which to install the stellar sensor, making sure that during the mission the stellar sensor is not turned towards the earth but that it is turned on the opposite side with respect to the earth. In the case of particular missions, it may be necessary to install more than one stellar sensor with a very significant burden both in terms of design (it is necessary to provide for additional dedicated connections for the additional stellar sensors) and in terms of costs. In this regard, the Applicant has in fact verified that the stellar sensor must be very sensitive (and therefore expensive) in order to discern also dimly lit stars, since it is not possible to guarantee a priori that the portion of stars framed within the limited field of view include bright stars.

Similarly, the Applicant has verified that since the angular diameter of the sun, seen from the vicinity of the earth, is about 0.5 degrees, the sun sensor must be carefully positioned to prevent the terrestrial albedo from masking or otherwise distorting the reading of the sun sensor. In any case, it is often necessary to correct the reading of the sun sensor in order to take into account the terrestrial albedo by resorting to mathematical models that try to describe the terrestrial albedo when the seasons vary and when the relative position changes with respect to the earth. These mathematical models, in addition to requiring a high computational cost from the AD system, are not always reliable and cannot take into account events that are not considered in the mathematical model but which are able to influence the real albedo. Even in the case of sun sensors it is therefore sometimes necessary to prepare and install more than one sensor on the satellite in order to be able to guarantee to determine the position of the sun, with an increase in terms of design and, at least in part, of costs.

Similar considerations also apply to the other types of sensors used in the AD system.

Therefore, in the Applicant's experience, each satellite requires a dedicated attitude determination system that is specifically designed and built to be able to operate on the specific satellite and for the specific mission.

The Applicant has perceived that by increasing the number of sensors installed on the satellite, the flexibility in positioning them increases and the information relating to the mission and the attitude that the satellite will have to assume during the mission decreases.

However, the Applicant has noted that this entails a greater expenditure in terms of dedicated connections which are necessary to ensure the correct operation of the sensors themselves, complicating the design of the AD system of the satellite and making the integration of the sensors in the satellite even more critical.

However, the Applicant has perceived that by making a heterogeneous group of sensors of the AD system independent and autonomous, it would be possible to operate this heterogeneous group of sensors as an independent satellite module requiring a single data connection and a single electricity connection for the correct operation of the whole module. In this way the design complexity for integrating the sensors in the satellite would be significantly reduced.

The Applicant has also perceived that by equipping said satellite module with at least two sensors of different type oriented differently from each other, i.e. oriented in different directions, it is possible to equip the satellite with more satellite modules identical to each other but oriented to each other differently so that each type of sensor can have, overall, an observation field of any angular diameter, even 360 degrees.

The Applicant also perceived that such a satellite module would considerably decrease the cost of the AD system. In fact, with even a single type of AD satellite module, it would be possible to equip any satellite, allowing a large-scale production of AD satellite modules that are identical between them. Even if different types of satellite modules are to be envisaged, each of them equipped with specific sensors, it would be possible to massively produce each type of AD satellite module and choose, from time to time, the satellite modules to be used.

The present invention therefore relates, in a first aspect, to a satellite module for attitude determination that includes a polyhedral containment body configured to be mounted on a satellite and comprising at least one data acquisition board and a connection interface for allowing the data acquisition board to be in signal communication and electrical communication with the satellite; at least one first-type sensor selected from a sun sensor, an earth sensor, a stellar sensor, a horizon sensor, in communication with the data acquisition board and at least one second-type sensor, different from the first type, selected from a sun sensor, an earth sensor, a stellar sensor, a horizon sensor, and in communication with the data acquisition board; wherein the connection interface is mounted on a first face of the polyhedral containment body, the first-type sensor is mounted on a second face of the polyhedral solid body, and the second-type sensor is mounted on a third face of the polyhedral solid body.

The present invention relates, in a second aspect, to a satellite comprising at least two satellite modules, preferably at least four satellite modules, even more preferably eight satellite modules, in accordance with the first aspect of the invention.

The polyhedral containment body allows the satellite module to arrange surfaces (faces), differently oriented between them, that can be used to mount sensors of different types which are then mounted on the containment body so as to be oriented to each other in a different way.

The data acquisition board allows collecting the electrical signals coming from the sensors. The data acquisition board can further comprise a circuit for conditioning the received signal to convert said signal into digital values, possibly making use of an analogue-digital converter. The data acquisition board may further comprise or be in signal communication with a microcontroller configured to process the digital signal and generate an output configured to be received by the satellite attitude control system. The microcontroller can be mounted in the containment body or can be mounted in the satellite outside the containment body.

The connection interface allows exchanging signals between the data acquisition board and the satellite attitude control system and allows energizing the data acquisition board and possibly the sensors (if the sensors need an electrical power supply).

The connection interface may be a single connection plug or may comprise multiple connection plugs. In any case, the connection interface is mounted on a face of the containment body that is different from the faces that house the sensors, so that when the module is connected with the satellite, the field of view of the sensors is not compromised.

The satellite module is therefore configured as an autonomous and independent module, that is to say as an independent satellite module that requires a single data connection and a single electricity connection for the correct operation of the sensors.

The term “polyhedral body” refers in the present description and in the following claims to a body defined by a finite number of polygons, in which each polygon defines a face, and in which the intersection of two faces is an edge or a vertex, each edge belonging only to two faces, two adjacent faces not being coplanar. Each polygon is preferably contained in a plane. Alternatively, at least one or all of the polygons may not be plane but they can be curved surfaces.

The term “satellite” refers in the present description and in the following claims to any vehicle capable of moving in space or in the upper atmosphere, including objects orbiting around a celestial body and spacecrafts capable of travelling in space or around an orbit.

The term “angular diameter” (ad), when referring to a field of view, in the present description and in the following claims means the measurement of the diameter (d) of the circle of smaller diameter capable of circumscribing an object entirely included in the field of view, with respect to the distance (D) from the observer, according to the formula ad=2*arctan (d/2*D).

The present invention, according to the first or second aspect, can comprise, individually or in combination, one or more of the following characteristics.

Preferably, the first-type sensor comprises an earth sensor or a sun sensor.

Preferably, the second-type sensor is a stellar sensor.

The Applicant has verified that by equipping a satellite with at least two identical satellite modules and by differently orienting the polyhedral containment body of one satellite module with respect to the other one, it is possible, depending on the shape of the satellite, to obtain fields of view for each type of sensor capable of ensuring a correct reading of the respective quantities regardless of the attitude assumed by the satellite.

In any case, the Applicant has verified that it is always possible, whatever the shape of the satellite, to choose a number of satellite modules that guarantee field of views for each type of sensor capable of ensuring a correct reading of the respective quantities regardless of the attitude assumed by the satellite.

By way of example, by placing eight satellite modules at the eight vertices of a cuboid-shaped satellite, it is possible to orient each module so that both the first-type sensors and the second-type sensors are oriented perpendicularly to the six faces of the cuboidal satellite, with also at least two sensors of each type oriented redundantly.

In this way, it is not necessary to know a priori the mission details and the attitude that the satellite will have to assume in order to correctly equip the satellite with an AD system.

The Applicant has also perceived that the satellite module according to the present invention allows to use stellar sensors having a very low cost.

In fact, the stellar sensors that are usually used have a very small field of view (FOV) (angular diameter between 4 degrees and 7 degrees) and therefore need to be very sensitive, and therefore expensive, in order to detect a sufficient number of stars or sufficiently bright stars to proceed with a comparison with a stellar catalogue.

The Applicant has noted that by mounting the satellite modules on the satellite in such a way that each stellar sensor is pointed in a different direction from the others, the probability that there is at least one stellar sensor not turned towards the earth or the sun is very high regardless of the attitude assumed by the satellite. By arranging eight satellite modules at the eight vertices of a satellite (for example a cubic shaped satellite), the probability that there is at least one stellar sensor not turned towards the earth or the sun is 100%.

Therefore, the Applicant has noted that the field of view of the stellar sensor can be greatly increased, since even in the face of a high field of view it is substantially certain that at least one stellar sensor is framing a portion of space that does not include the earth or the sun.

The Applicant has perceived that by increasing the field of view of the stellar sensor it is possible to decrease the sensitivity thereof (and therefore decrease the cost of the stellar sensor), since it is very likely that with a wide field of view a sufficient number of very bright stars will always be framed so as to proceed with the comparison with a stellar catalogue.

The Applicant has also perceived that by using a stellar sensor with a wide field of view and low sensitivity it is also possible to reduce the energy expenditure of the AD system.

In fact, the portion of stars framed by a stellar sensor with a reduced field of view must be compared with very accurate stellar catalogues in order to be sure of being able to recognise the small portion of stars framed. Typically, a stellar sensor with a field of view having an angular diameter of about 4 degrees requires a stellar catalogue of hundreds of thousands of stars. The computational cost for comparing the portion of stars framed with said stellar catalogue is very expensive and requires a considerable amount of electricity, up to a few tens of Watts, to power the processors implemented for the comparison.

On the contrary, the Applicant has noted that the portion of stars framed by a stellar sensor with a wide field of view can be compared with much less accurate stellar catalogues. The Applicant has verified that a stellar catalogue containing a few hundred stars is sufficient, for example about 500 stars. Therefore, the computational cost for the comparison is very low and requires much less energy.

Preferably, the stellar sensor has a field of view having an angular diameter higher than 15 degrees.

Even more preferably, the stellar sensor has a field of view having an angular diameter between about 20 degrees and about 60 degrees, preferably of about 40 degrees.

Preferably, the first-type sensor has a field of view having an angular diameter of at least 60 degrees.

Preferably, the satellite module comprises a third-type sensor, different from the first and second types, mounted on a fourth face of the polyhedral containment body.

The third-type sensor increases the accuracy in attitude determination as it provides additional data, based on a different measured quantity, which can be processed to determine the attitude.